1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically to turbine rotor blade with blade tip cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as a large frame heavy duty industrial gas turbine (IGT) engine, includes a turbine with one or more rows of stator vanes and rotor blades that react with a hot gas stream from a combustor to produce mechanical work. The stator vanes guide the hot gas stream into the adjacent and downstream row of rotor blades. The first stage vanes and blades are exposed to the highest gas stream temperatures and therefore require the most amount of cooling.
The efficiency of the engine can be increased by using a higher turbine inlet temperature. However, increasing the temperature requires better cooling of the airfoils or improved materials that can withstand these higher temperatures. Turbine airfoils (vanes and blades) are cooled using a combination of convection and impingement cooling within the airfoils and film cooling on the external airfoil surfaces.
A prior art turbine rotor blade cooling circuit is shown in FIG. 1 and includes a (1+3) serpentine flow cooling circuit for a first stage turbine rotor blade, the first stage being exposed to the highest gas stream temperatures in the turbine and therefore requires the most cooling. In the FIG. 1 prior art blade, the airfoil leading edge is cooled with a backside impingement cooling in conjunction with a leading edge showerhead arrangement of film cooling holes 11 along with pressure side and suction side gills holes 12. The cooling air for the leading edge region cooling is supplied through a separate radial cooling air supply channel 13 in which the cooling air passes through a row of metering and impingement holes 14 to produce backside impingement cooling of the leading edge region wall. The spent impingement cooling air is collected in the leading edge impingement channel or cavity 15 prior to being discharged through the showerhead film cooling holes and gills holes 12. The airfoil main body is cooled with a triple pass forward flowing (toward the leading edge) serpentine flow cooling circuit with a first leg 21 being the cooling air supply channel, a second leg 22 and a third leg 23 located adjacent to the leading edge cooling supply channel 13 along with pressure side film cooling holes and suction side film cooling holes. A row of trailing edge exit slots or holes discharge cooling air from the first leg out through the trailing edge region. FIG. 2 shows a flow diagram for the FIG. 1 blade cooling circuit and includes the blade tip cooling holes connected to the radial channels 13, 15, 21, 22 and 23 to provide cooling for the blade tip. FIG. 3 shows a cross section side view of the blade cooling circuit of FIG. 1.
For the blade of FIG. 1, the blade tip cooling is accomplished by drilling holes into the upper extremes of the serpentine flow cooling passages from both of the pressure and suction side surfaces near to the blade tip edge and the top surfaces of the squealer cavity. Film cooling holes are formed along the airfoil pressure side and the suction side tip sections. In addition, convection cooling holes formed along the tip rail at an inner portion of the squealer pocket provide for additional cooling for the squealer tip rail. Since the blade tip region is subject to severe secondary flow fields, this requires a large quality of film cooling holes and cooling air flow to provide adequate cooling for the blade tip periphery. FIG. 4 shows the FIG. 1 blade tip section from the pressure side wall with the row of tip periphery film cooling holes with FIG. 5 showing the tip peripheral film cooling holes for the suction side wall.
For the blade cooling circuit of FIG. 1, the last leg of the serpentine flow cooling circuit is predetermined by a ceramic core manufacturing requirement. As a result of the cooling design requirement, when the cooling air is bled off from the cavity for the cooling of both the pressure and suction sidewalls as well as along the blade tip section, the spanwise internal Mach number (velocity of the cooling air flow) decreases. This decreasing of the Mach number results in a lower through-flow velocity and cooling side internal heat transfer coefficient. This same decreasing Mach number with lower through-flow velocity and cooling side internal heat transfer coefficient also occurs in the airfoil leading edge cooling supply channel 13.